A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a rotor and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the rotor and disc assembly, to rotate.
Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades comprise a root, a platform and an airfoil that extends outwardly from the platform. The airfoil ordinarily comprises a tip, leading edge and a trailing edge. Most blades typically contain internal cooling channels forming a cooling system. The cooling channels in the blades may receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade.
Operation of a turbine engine results in high stresses being generated in numerous areas of a turbine blade. One particular area of high stress is found in the airfoil trailing edge, which is a portion of the airfoil forming a relatively thin edge that is generally orthogonal to the flow of gases past the blade and is on the downstream side of the airfoil. Because the trailing edge is relatively thin and an area prone to development of high stresses during operation, the trailing edge is highly susceptible to formation of cracks which may lead to failure of the airfoil.
A conventional cooling system in the airfoil of a turbine blade assembly may include cooling fluid passages to maximize convection cooling in the airfoil trailing edge, and discharge a substantial portion of the cooling air through the trailing edge of the airfoil. For example, a typical trailing edge cooling configuration comprises providing trailing edge cooling holes, which are conventionally of a constant diameter and are fed from a common cooling supply cavity, and which discharge at the centerline of the airfoil trailing edge or exit at an angle on the pressure side adjacent to the trailing edge. In the described arrangement, the cooling flow distribution into the trailing edge cooling holes and the pressure ratio across the cooling holes is predetermined by the cooing air pressure in the cooling supply cavity. The cooling air passing through the cooling holes is subsequently injected into the mainstream of hot gases, and may cause turbulence, coolant dilution and, in the case where pressure side bleed cooling is employed, there may be a loss of downstream film cooling effectiveness.
While many of the conventional airfoil cooling systems have operated successfully, a need still exists to provide increased cooling capability in the trailing edge portions of turbine blade airfoils while minimizing or reducing the flow of coolant into the mainstream gas flow.